XFOIL Version 6.94 Calculated polar for: WBL FX 71-089A AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0047 0.02862 0.01602 -0.0114 0.9989 0.3024 -2.750 0.0328 0.02861 0.01622 -0.0112 0.9989 0.3358 -2.500 0.0605 0.02864 0.01640 -0.0111 0.9989 0.3644 -2.250 0.0873 0.02861 0.01667 -0.0109 0.9989 0.3974 -2.000 0.1136 0.02861 0.01696 -0.0106 0.9989 0.4293 -1.750 0.1398 0.02865 0.01725 -0.0104 0.9989 0.4518 -1.500 0.1648 0.02864 0.01768 -0.0100 0.9989 0.4901 -1.250 0.1893 0.02863 0.01819 -0.0095 0.9989 0.5322 -1.000 0.2125 0.02850 0.01880 -0.0087 0.9989 0.5906 -0.750 0.2311 0.02798 0.01946 -0.0063 0.9989 0.7139 -0.250 0.3421 0.02824 0.01743 -0.0100 0.4533 1.0011 0.000 0.3713 0.02962 0.01812 -0.0100 0.4042 1.0011 0.250 0.4013 0.03088 0.01890 -0.0102 0.3715 1.0011 0.500 0.4322 0.03214 0.01985 -0.0105 0.3478 1.0011 0.750 0.4631 0.03346 0.02093 -0.0109 0.3273 1.0011 1.000 0.4942 0.03493 0.02231 -0.0113 0.3132 1.0011 1.250 0.5251 0.03670 0.02388 -0.0117 0.3029 1.0011 1.500 0.5572 0.03840 0.02579 -0.0125 0.2941 1.0011 1.750 0.5867 0.04045 0.02765 -0.0129 0.2847 1.0011 2.000 0.6176 0.04227 0.02979 -0.0138 0.2744 1.0011 2.250 0.6473 0.04455 0.03205 -0.0144 0.2683 1.0011 2.500 0.6776 0.04732 0.03505 -0.0155 0.2650 1.0011 2.750 0.7082 0.05030 0.03849 -0.0169 0.2631 1.0011 3.000 0.7370 0.05347 0.04211 -0.0185 0.2584 1.0011 3.250 0.7637 0.05655 0.04540 -0.0195 0.2524 1.0011 3.500 0.7900 0.06065 0.04988 -0.0212 0.2516 1.0011 3.750 0.8134 0.06461 0.05389 -0.0218 0.2469 1.0011 4.000 0.8347 0.06960 0.05934 -0.0241 0.2463 1.0011 4.250 0.8504 0.07819 0.06870 -0.0298 0.2645 1.0011 4.750 0.6509 0.12089 0.11249 -0.0790 0.6030 1.0011 5.000 0.6691 0.12421 0.11578 -0.0777 0.5693 1.0011